Appendix 9 Section 23.1323(b) requires the system error, including position error, but excluding instrument error, not to exceed 3 percent of CAS or 5 knots, whichever is greater, in the designated speed range.
(3) Compressibility. For many years CAS was used for design airspeeds. However, as speeds and altitudes increased, a compressibility correction became necessary because airflow produces a total pressure on the pitot head that is greater than if the flow were incompressible. We now use EAS as a basis for design airspeeds (§ 23.235). Values of CAS versus EAS may be calculated or you may use the chart in Appendix 7, figure A7-5, to convert knots CAS to EAS.
3. GPS Method. The Global Positioning System (GPS) method consists of using a GPS unit to determine ground speed, which is then used to calculate true airspeed. Any commercial GPS unit can be used that produces consistent results. Once true airspeed is calculated, the data reduction is nearly identical to the speed course method described previously. One difference is that the scale altitude correction factor (ΔVc the difference between CAS and EAS as shown in figure A7-5) may be significant at higher altitudes and speeds that may be flown with this method. Specifically, you will solve the following equation for (ΔVpec):
Vi + ΔVic + ΔVpec + ΔVc = Vtrue* And then, assuming that all of the pitot-static error is in the static port, you may calculate the altitude position, error, ΔHpec, as described in the Trailing Bomb/Cone Method.
a. Test Conditions.
(1) Air Quality. The air should be as smooth as possible with a minimum of turbulence. The wind velocity and direction must be constant for this method to give correct results.
(2) Weight and C.G. Same as the speed course method.
(3) Altitude. The altitude is not critical, but it should be chosen where the air is smooth and the winds are constant.
(4) Speed Range. Any speed at which the aircraft can be stabilized in level flight (see Figure A9
2).
(5) Runs . Three runs per airspeed are required to calculate one true airspeed. The three runs must be done at the same indicated speed and altitude on different headings. The headings should be 60 to 120 degrees apart. _____
(6) Configuration. Same as the speed course method.
A9-12
Appendix 9
b. Test Procedures.
(1) Stabilize the airplane in steady level flight at the desired test speed configuration. Record the indicated airspeed, pressure altitude (29.92 set), outside air temperature and configuration of the aircraft.
(a) Record both ground track and ground speed from the GPS unit once the values are stable (this can take up to 10 seconds after stabilizing).
(b) Turn 60 to 120 degrees either direction and record the new ground track and speed once restabilized at the same airspeed and altitude on the new heading. Minor variations in altitude (up to 100 feet) are much preferred to any variation in airspeed from the initial value. A one knot change in indicated airspeed will cause at least a one knot change in true airspeed, but 100 feet of altitude only causes on the order of 1/2 of 1 percent change in the density ratio, σ.
(c) Turn another 60 to 120 degrees in the same direction and record a third set of ground track and speeds.
(2) Once you have three sets of ground track and speed for a given indicated airspeed and configuration, repeat steps (1)(a) through (1)(c) above at sufficient increments, to provide an adequate calibration curve for each of the configurations.
A9-13
Appendix 9
c. Data Reduction. The best way to calculate true airspeed from the three sets of ground tracks and speeds is with a personal computer spreadsheet. The following solution was developed assuming that the three legs flown had the same true airspeed (indicated airspeed, OAT, and pressure altitude were identical) and that the wind did not change during the three legs. The table shows the spreadsheet equations for one popular spreadsheet program that will solve the problem. Note that wind speed and direction are intermediate outputs. If a series of points are done at nearly the same time, altitude and geographic location, then the consistency of the calculated wind speed and direction will be an indicator of the validity and accuracy of the calculated true airspeeds.
A B Result 1 Ground Speed 1 184 184 2 Track 1 265 265 3 Ground Speed 2 178 178 4 Track 2 178 178 5 Ground Speed 3 185 185 6 Track 3 82 82 7 X1 =B1*SIN(PI()*(360-B2)/180) 183.3 8 Y1 =B1*COS(PI()*(360-B2)/180) -16.0 9 X2 =B3*SIN(PI()*(360-B4)/180) -6.2 10 Y2 =B3*COS(PI()*360-B4)/180) -177.9 11 X3 =B5*SIN(PI()*(360-B6)/180) -183.2 12 Y3 =B5*COS(PI()*(360-B6)/180) 25.7 13 M1 =-1*(B9-B7)/(B10-B8) -1.17 14 B1 =(B8+B10)/2-B13*(B7+B9)/2 6.71 15 M2 =-1*(B11-B7)/(B12-B8) 8.77 16 B2 =(B8+B12)/2-B15*(B7+B11)/2 4.42 17 Wx =(B14-B16)/(B15-B13) 0.2 18 Wy =B13*B17+B14 6.4 19 Wind Speed =SQRT(B17^2+B18^2) 6.4 20 Wind Direction =MOD(540-(180/PI()*ATAN2(B18,B17)),360) 177.9 21 True Airspeed =SQRT((B7-B17)^2+(B8-B18)^2) 184.4
4. Pace Airplane Method. An airplane whose pitot static systems have been calibrated by an acceptable flight test method is used to calibrate the pitot static systems of a test aircraft.
a. Test Conditions. Smooth ambient flight conditions
b. Test Procedures. The pace airplane is flown in formation with the test airplane at the same altitude and speed. The aircraft must be close enough to ensure that the relative velocity is zero yet far enough away so that the pressure fields of the two airplanes do not interact. Readings are coordinated by radio.
c. Data to be recorded
(1) Test Airplane airspeed (ViT) kts
(2) Test Airplane Pressure Altitude (HiT) ft A9-14
Appendix 9
(3) Pace Airplane airspeed (Vip) kts
(4) Pace Airplane Pressure Altitude (Hip) ft
(5) Configuration for both airplanes
(6) Fuel used in both airplanes
d. Data Reduction. Correct all the instrument readings for instrument errors and the pace aircraft readings for the known position error.
VpecT (Vip Vicp Vpec ) (ViT VicT )kts H pecT (Hip H icp H pec ) (HiT HicT ) ft Calculate VpecT and H pecT for all data points in each configuration and plot in a manner similar to Figure A9-4 and Figure A9-5.
5. Pitot-Static Boom Data. If a flight test Pitot-Static boom is mounted on an airplane such that the pitot tube (total head) is not affected by flow angularity and the static source is outside the pressure field of the aircraft, then it can be assumed that the boom data is without position errors. The boom data can then be taken as the pace data. The Pitot-Static boom should be calibrated for high alpha and high speed.
a. Test Conditions. Smooth ambient flight conditions
b. Test Procedures. The nose or wing-mounted boom is flown in formation with the test airplane at the same altitude and speed. The aircraft must be close enough to ensure that the relative velocity is zero yet far enough away so that the pressure fields of the two airplanes do not interact. Readings are coordinated by radio.
c. Data to be recorded
(1) Test Airplane airspeed (ViT) kts
(2) Test Airplane Pressure Altitude (HiT) ft
(3) Nose or wing-mounted boom airspeed (Vip) kts
(4) Nose or wing-mounted boom Pressure Altitude (Hip) ft
(5) Configuration for both airplanes
(6) Fuel used in both airplanes A9-15
Appendix 9
d. Data Reduction.
Vpect (ViB VicB Vpec ) (ViT VicT ) kts H pect (H iB H icB Vpec ) (H iT HicT ) kts ft VpecT and H pecT are calculated throughout the speed range in each configuration and plotted as shown in figures 3 and 4.
6. Tower Fly-by Method. The tower fly-by method is one of the methods which results in a direct determination of static error in indicated pressure altitude.
Since the altimeter and airspeed system use the same static source, it is possible to correlate the altimeter position error directly to the airspeed error. This correlation assumes that there is no error in the total head system.
The tower fly-by method can be modified to use radar altimeter or differential GPS to determine geometric/tapeline height above a ground based pressure measuring station.
Figure A9-6 Tower Fly-By Method
a. Procedures and Test Conditions for Tower Fly-by
(1) Air Quality. Smooth, stable air is needed for determining the error in pressure altitude.
(2) Weight and C.G. Same as for calibrations of the airspeed indicating system.
(3) Speed Range. The calibration should range from 1.3 VSO to 1.8 VS1. Higher speeds up to VMO or VNE are usually investigated so that errors can be included in the AFM for a full range of airspeeds.
A9-16
-- Appendix 9
(4) Test Procedures.
(i) The test technique is to fly the aircraft along a ground reference line, past the tower, in stabilized flight at a constant airspeed and at the approximate height of the tower. The primary piloting task is to maintain a constant indicated altitude during the run. The tower is equipped with a sensitive altimeter and a means of determining the relative angle ( ) of the aircraft. The data recorded during each run are the indicated pressure altitude of the tower, (Hitower), the angle ( ), and the aircraft's indicated pressure altitude, airspeed and temperature (HiA/C, ViA/C, and TiA/C) as it passes the tower. Note that the tower altimeter must be at the zero grid line position in the tower.
(ii) Repeat step (i) at various airspeeds in increments sufficient to cover the required range at each flap setting.
(5) Data Acquisition. Data to be recorded at each test point:
(i) Airplane Airspeed ViA/C knots.
(ii) Airplane indicated pressure attitude, HiA/C knots.
(iii) Tower observer indicated pressure altitude, Hitower.
(iv) Angle of aircraft above the tower.
(v) Wing flap position.
(vi) Landing gear position.
(vii) Fuel used in airplane.
(viii) TiA/C and Titower
(6) Data Reduction. The actual pressure altitude of the aircraft is Hcref where TsHcref = (Hitower + Hictower) + D tan Tt Where Ts is the standard day absolute temperature at the test altitude and Tt is the test day temperature in absolute units.
A9-17
Appendix 9 TThe s temperature correction is to convert the geometric height of the aircraft above the Tt reference zero grid line in the tower (D tan ) to a pressure height that can be added to the pressure altitude of the tower Hctower. The difference between the actual refere nce pressure altitude of the aircraft and the aircraft's instrument-corrected pressure altitude is the position error correction.
Hpec = Hcref - (HiA/C + HicA/C) Ts = [(Hitower + Hictower) + D tan ] - (HiA/C + HicA/C)Tt Hpec is calculated for every speed and aircraft configuration flown past the tower and the data are plotted as per Figure A9-4.
The airspeed system position error corrections can be obtained from the tower fly-by method if it is assumed that the pitot tube (total head) errors are zero.
The hydrostatic equilibrium equation states that the pressure error correction at the static source is ps = - g Hpc and from Section 3.
2 4 2 41 2 M c M c 1 2 M ic M icpd pp ps oVc (1 ....) oVic (1 ....)2 4 40 2 4 40 Since it is assumed that p 0 and for low speed aircraft, compressibility effects can be 1 2 2ignored, then p p (V V )d s o c ic 2 The above equation is used to calculate VC at every test point, then Vpec Vc Vic . The data are then plotted as per Figure A9-5.
7. Ground Run Airspeed System Calibration. The airspeed system is calibrated to show compliance with commuter category requirements of § 23.1323(c) during the accelerate-takeoff ground run, and is used to determine IAS values for various V1 and VR speeds. The airspeed system error during the accelerate-takeoff ground run may be determined using a trapped static source reference or a distance measuring unit that provides readouts of ground speed that can be converted into CAS.
a. Definitions.
(1) Ground Run System Error. System error during the accelerate-takeoff ground run is the combination of position error, instrument error, and the dynamic effects, such as lag, which may be caused by acceleration on the runway.
(2) Trapped Static Source. An airtight bottle with sufficient internal volume so as to be infinite when compared to an airspeed indicator's internal changes in volume while sensing various airspeeds. The bottle should be insulated to mi nimize internal bottle temperature changes as testing is in progress. For short periods of time, it can be assumed that the bottle will reflect true static ambient pressure to the test indicator.
A9-18
Appendix 9
(3) Production Airspeed Indicator. An airspeed indicator that conforms to the type certification design standards. The indicator should be installed in the approved instrument panel location since these tests involve the dynamic effects of the indicator that may result from acceleration.
(4) Test Airspeed Indicator. A mechanical airspeed indicator with known dynamic characteristics during acceleration or an electronic transducer that can provide airspeed information.
(5) Test Reference Altimeter. An altimeter that indicates the altitude of the air trapped in the bottle or local ambient static air if the valve is opened.
(6) Ground Run Position Error. Ground run position error is the static-pressure error of the production static source during ground runs with any ground effects included. Any contributions to error due to the total-pressure (pitot) are ignored.
(7) Instrument Error. See paragraph 302a(3)(b).
(8) Dynamic Effects on Airspeed Indicator. The dynamic effects on airspeed indicators occur as a result of acceleration and rapid change in airspeed during takeoff. This causes many airspeed indicators to indicate an airspeed lower than the actual airspeed.
Note: It is possible for electronic airspeed indicators driven by an air data computer to also have errors due to dynamic acceleration effects because of characteristics inherent in the basic design.
(9) Distance Measuring Unit. An instrumentation system normally used to record takeoff distance measurements. One output of these systems provides the ground speed versus time as the airplane accelerates during the accelerate-takeoff ground run. Ground speed may be converted into a corresponding CAS value by applying wind and air density corrections at intervals during acceleration where the ship's airspeed indications have been recorded.
b. Trapped Static Source Method. The trapped static source method consists of comparing instantaneous readings of airspeed, as indicated on a test airspeed indicator, with readings on a production airspeed indicator while accelerating on each other. Readings may be recorded by film or video cameras for mechanical airspeed indicators or by electronic means if a transducer type device is being utilized.
See Figure A9-7 for system schematic.
(1) Test Conditions.
(a) Air Quality. The surface winds should be light with a minimum of gusting.
(b) Weight and C.G. Ground run calibrations are not sens itive to c.g. The dynamic effects of acceleration may be affected by weight. Test weight variations should be sufficient to account for any measurable effects due to weight.
(c) Speed Range. The speeds should range from 0.8 of the minimum V1 to 1.2 times the maximum V1, unless higher values up to VR are required for expansion of takeoff data.
A9-19
Appendix 9
(d) Configuration. The airspeed system should be calibrated during the accelerate-takeoff ground run for each approved takeoff flap setting.
(2) Test Procedures.
(a) Align the airplane with the runway.
(b) With idle engine power, and with the cabin door open, open the valve to expose the bottle to static ambient conditions, then close the valve. Record the test altimeter reading.
(c) Close the cabin door.
(d) Conduct a takeoff acceleration using normal takeoff procedures. The camera should be recording speeds from the two airspeed indicators in increments sufficient to cover the required airspeed range.
A9-20
Appendix 9 Figure A9-7 - Trapped Static Source Schematic
TO SHIP'S TO SHIP'S
PITOT PITOT
SOURCE SOURCE
PRODUCTION TEST AIRSPEED
SHIP'S INDICATOR
AIRSPEED (OR
INDICATOR ELECTRONIC DEVICE, I.E.
TRANSDUCER) TEST
REFERENCE
ALTIMETER
PRODUCTION OPENABLE OPEN TO
RATE-OF- VALVE AMBIENT STATIC
CLIMB AIR CONDITIONS
TO SHIP'S
STATIC
SOURCE
(e) The takeoff run should be continued, if possible, until beyond the maximum required speed then aborted. When at rest with engi nes idling, open valve again and observe the test altimeter. Any significant jumps or changes in indicated altitude may indicate a system leak, too much runway gradient or other factors WHICH WILL INVALIDATE THE RESULTS OF THE RUN.
(f) Repeat steps (a) through (e) of this paragraph until there are sufficient runs to provide adequate calibration curves for the required configurations.
A9-21
Tfi)IE TRAP PED
(1) TS CORREC TED SHIP 1SIAS
(l)SHIP 1S
CORRECTED
(2) AIRS PEED STATIC (TS) IAS AIRSP EED TSIA S (KN OTS)
AIRSPEED SHIP 'SIAS ERRO R (KNOTS) INS
TRUMENT INST RUM ENT
CORRECTION
CORRECTIO N 7:41 :45
50.7 0
50.7 49 0 49
1.7 :46
56.1 I
56.1 54 I 54 +2.1 :47
61.4 I
61.4 61 I 61 +o.4 :48
66.9 I
66.9 66 I 66 +-0.9 :49
71.9 I 7l.9 72 I 72 -0.1 50
76.7 I
76.7 77 I 77 -0.3 :51
82.1 I
82.1 83 I 83 -0.9 :52
86.8 I
86.8 88 I 88 -l.2 :53
91.5 I
91.5 91 I 91 +o.5 :54
96.5 I
96.5 99 I 99 -2.5 :55
100.9 I I00.9 102 I !02 -I.I :56 !05.2 I !05.2 !07 I 107 -1.8 :57
110.1 I IIO.l 113 I IIJ -2.9 :58
114.4 I
114.4 119 I ll9 -4.6 :59
118.2 I
118.2 123 I 123 -4.8 7:42 :00
122.9 V
122.9 128 V 128 -5.1 11/16/2011 AC 23-8C Appendix 9
(3) Data Acquisition and Reduction. Read the recorded data (film or video) at increments of airspeed arbitrarily selected within the required range. See Table A9-2 for a sample data reduction. Record and perform the following:
Table A9-2 - Trapped Static (TS) Data Reduction NOTES: 1. Obtain from instrument calibration.
2.
Corrected trapped static IAS minus corrected ship's IAS.
3.
Corrections are added.
A9-22